1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to a ring segment for a turbine in an industrial gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
A row or stage of turbine rotor blades rotate within an annular arrangement of ring segments in which blade tips form a small gap with an inner or hot surface of each ring segment. The size of the gap changes due to different thermal properties of the blade and the ring segments from a cold sate to a hot state of the turbine. The smaller the gap, the less hot gas leakage will flow between the blade tips and the ring segments.
An IGT engine operates for long periods of time at steady state conditions, as opposed to an aero gas turbine engine that operates for only a few hours before shutting down. Thus, the parts in the IGT engine must be designed for normal operation for these long periods, such as up to 40,000 hours of operation at steady state conditions. A thin TBC (Thermal Barrier Coating) is applied to the inner or hot surface of each ring segment in order to insulate the ring segment from the hot gas flow and reduce the metal temperature of the ring segment. A reduced metal temperature requires less cooling air flow and thus improves the turbine efficiency. As the turbine inlet temperature increases, the cooling flow demand for cooling the ring segments will also increase and therefore reduce the turbine efficiency. One method of reducing the cooling air consumption while allowing for higher turbine inlet temperatures is to use a thicker TBC and film cooling for the ring segments. Thus, the design of the cooling circuit for the ring segments relies more on the endurance of the TBC. Therefore, the TBC becomes the main factor in the design of the ring segment cooling circuit. A problem is that the thicker the TBC the higher the chance of spallation (when pieces of the TBC break away).